Turbine engine nozzle

ABSTRACT

A turbine engine nozzle assembly has an upstream flap assembly having a main flap and a liner, a cooling passageway formed between the main flap and liner. A downstream flap is pivotally coupled to the upstream flap assembly for relative rotation about a hinge axis. The liner has a trailing end spaced upstream from a trailing end of the main flap by at least 40% of a length of the main flap.

U.S. GOVERNMENT RIGHTS

The invention was made with U.S. Government support under contract no.N00019-02-C-3003 awarded by the U.S. Navy. The U.S. Government hascertain rights in the invention.

BACKGROUND OF THE INVENTION

The invention relates to turbine engines. More particularly, theinvention relates to variable throat turbine engine exhaust nozzles.

There is well developed field in turbine engine exhaust nozzles. Anumber of nozzle configurations involve pairs of relatively hingedflaps: a convergent flap upstream; and a divergent flap downstream.Axisymmetric nozzles may feature a circular array of such flap pairs.Exemplary nozzles are shown in U.S. Pat. Nos. 3,730,436, 5,797,544, and6,398,129 and United Kingdom patent application GB2404222 A.

SUMMARY OF THE INVENTION

Accordingly, one aspect of the invention involves a turbine enginenozzle subassembly. An upstream flap assembly has a main flap and aliner. A cooling passageway formed between the main flap and liner. Adownstream flap is pivotally coupled to the upstream flap assembly forrelative rotation about a hinge axis. The liner has a trailing endspaced upstream from a trailing end of the main flap by at least 40% ofa length of the main flap.

The details of one or more embodiments of the invention are set forth inthe accompanying drawings and the description below. Other features,objects, and advantages of the invention will be apparent from thedescription and drawings, and from the claims.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a cutaway longitudinal view of a prior art turbine enginenozzle.

FIG. 2 is a partial view of assembled flaps of the nozzle of FIG. 1

FIG. 3 is cutaway longitudinal view of a modified turbine engine nozzle.

FIG. 4 is an enlarged view of a convergent flap of the nozzle of FIG. 3.

FIG. 5 is a partial view of assembled flaps of the nozzle of FIG. 3.

FIG. 6 is an exploded view of a convergent flap liner member of thenozzle of FIG. 3.

FIG. 7 is an oblique outboard view of the liner member of FIG. 6.

FIG. 8 is an oblique inboard view of the liner member of FIG. 6.

FIG. 9 is an oblique inboard view of a convergent seal liner member ofthe nozzle of FIG. 3

FIG. 10 is an oblique outboard view of the convergent seal liner numberof FIG. 9.

FIG. 11 is a transverse sectional view of the nozzle of FIG. 3.

FIG. 12 is a schematic longitudinal sectional view of the nozzle of FIG.3.

FIG. 13 is a graph of gas mass flux for cooling and main flows againstliner length.

FIG. 14 is a graph relating blowing ratios of baseline and modifiednozzles.

Like reference numbers and designations in the various drawings indicatelike elements.

DETAILED DESCRIPTION

FIG. 1 shows a prior art turbine engine nozzle 20. The exemplary nozzleis drawn from United Kingdom patent application GB2404222 A (thedisclosure of which is incorporated by reference herein as if set forthat length) and may serve as a baseline for modifications discussedbelow. The exemplary nozzle 20 comprises an axisymmetric circular arrayof convergent/divergent flap pairs about a nozzle axis or centerline500. A given flap pair has a convergent flap 22 upstream/forwardextending from an upstream end 23 to a downstream end 24 and a divergentflap 26 downstream/aft extending from an upstream end 27 to a downstreamend 28. The flaps are hinged relative to each other by a hinge mechanism30 for relative movement about a hinge axis 502 proximate the convergentflap downstream end and divergent flap upstream end. A concentriccircular array of seals may be interspersed with the flap pair array.Exemplary seals (FIG. 2) have respective convergent members 32 anddivergent members 33 between adjacent convergent flaps and divergentflaps, respectively. The inboard surface of the divergent flap 26 has alongitudinally convex surface portion 40 (FIG. 1) near its upstream endfor forming an aerodynamic throat (i.e., the location of smallestpassageway cross-section) of the nozzle of instantaneous throat radiusR_(T) and an essentially longitudinally straight portion 42 extendingaft therefrom toward the downstream end for forming an exhaust outlet ofinstantaneous outlet radius R_(O). For each convergent/divergent flappair, the nozzle further includes an external flap 50, the outboardsurface 52 of which forms an exterior contour of the nozzle exposed toexternal airflow passing around the aircraft fuselage.

FIG. 1 further shows a nozzle static ring structure 60 for mounting thenozzle to the engine, aircraft fuselage, or other environmentalstructure. Proximate the upstream end 23 of the convergent flap 22, ahinge structure pivotally couples the convergent flap to the static ringstructure 60 for relative rotation about a fixed transverse axis 503. Asynchronization ring 62 is mounted between inboard and outboard aftportions 64 and 66 of the static ring structure and may belongitudinally reciprocated by actuators (e.g., pneumatic or hydraulicactuators-not shown). In the condition of FIG. 1, the synchronizationring is at a forwardmost/upstreammost position. The synchronization ringis coupled to each flap pair by an associated linkage 70. Each linkage70 includes a central bell crank 72 pivotally coupled by a hingemechanism to a bell crank ground point 74 at the trailing edge of thestatic ring structure inboard portion 64 for relative rotation about afixed transverse axis 504. To drive rotation of the bell crank throughits range of rotation about the axis 504, the bell crank is coupled tothe synchronization ring by an associated H-link 76. A forward end ofthe H-link is pivotally coupled to the synchronization ring by a hingemechanism for relative rotation about a transverse axis 506 which shiftslongitudinally with the synchronization ring. An aft end of the H-linkis pivotally coupled to the bell crank by a hinge mechanism for relativerotation about a transverse axis 508 which moves along a circular pathsegment centered about the axis 504 in response to linear translation ofthe axis 506. Thus, as viewed in FIG. 1, a rearward shift of thesynchronization ring produces a clockwise rotation of the bell crankabout the axis 504. Rotation of the bell crank is transferred toarticulation of the associated flap pair by an associated pair oftransfer links 78. Forward/upstream ends of each pair of transfer linksare pivotally coupled to the bell crank for relative rotation about atransverse axis 510 which also moves along a circular path segmentcentered about the axis 504 in response to linear translation of theaxis 506. Aft/downstream ends of the transfer links are pivotallycoupled to the divergent flap 26 for relative rotation about atransverse axis 512. As discussed below, in the exemplary embodimentmovement of the axis 512 is not entirely dictated by the rotation of thebell crank and associated static ring translation. Rather, it may beinfluenced by other forces, namely aerodynamic forces arising fromrelative pressures internal and external to the nozzle. In exemplaryembodiments, the axis 512 falls aft of the axis 502 and along a forwardhalf of the span between upstream and downstream ends of the divergentflap. More narrowly, it falls along a forward third, and, in theillustrated embodiment, approximately in between about the first 5% and15% of such span.

In the exemplary embodiment, the external flap 50 has a forward end 90pivotally coupled by a hinge mechanism to the static structure outboardportion 66 for relative rotation about a fixed transverse axis 520.Proximate its downstream end 92, the external flap is pivotally coupledby a hinge mechanism to the divergent flap 26 (slightly more forward ofits downstream end 28) for relative rotation about a transverse axis522. The external flap is configured so that the span between the axes520 and 522 is extensible and contractible such as by having an upstreamlink 94 telescopically mounted relative to a main body portion 96 of theexternal flap and coupling the external flap to the static ringstructure. The extensibility/contractability may have a limited range.For a further limitation on that range, a secondary link or mode strut100 is provided having a forward end portion 102 pivotally coupled tothe static ring structure for relative rotation about a fixed transverseaxis 524 which may be close to the axis 520. If the axes 520 and 524 arecoincident, it may be advantageous to drill one hole through all pivotpoints for low cost. However, if the width of the external flap 50 issuch that the main body portion 96 on either circumferential side of theflap are substantially circumferentially spaced from the mode strut, itmay be advantageous to locate the axis 520 relatively closer to theengine centerline than the axis 524 so as to maintain a good mechanicaladvantage for the mode strut.

An aft end portion 104 of the mode strut is pivotally coupled to thedivergent flap 26 for relative rotation about an axis 526 fixed relativeto the mode strut but floating relative to the divergent flap with arestricted range of movement. The exemplary range of movement isprovided by the use of a pair of mounting brackets 110 at anintermediate location on the divergent flap, each having a slot 112accommodating an obround slider 113 on a pivot shaft 114 fixed along theaxis 526 relative to the mode strut. The slider and shaft are free tomove along the slot between first and second ends 116 and 118 thereof.An exemplary intermediate location is, approximately within the middlethird of the divergent flap length and the middle third of the spanbetween axes 512 and 522.

In operation, the position of the synchronization ring 62 determines anominal throat radius R_(T) and associated throat area (i.e., a throttlecondition). In a given synchronization ring position, the aerodynamicforces may then determine the mode which is nominally associated withthe divergent flap interior surface angle θ. FIG. 1 shows thesynchronization ring at the forward extremity of its range of motion,thereby establishing the maximum nominal throat area. FIG. 1 furthershows a high mode condition in which the aerodynamic forces place thedivergent flap in its maximum θ condition with the slider 113 bottomedagainst the slot end 116. Under changed conditions, the force balanceacross the combination of external flap 50 and divergent flap 26 mayproduce an alternate θ. For example, in a maximum area, minimum θ lowmode condition the slider may be 113 is substantially bottomed againstthe slot end 118. In an alternate configuration, the operation of themode strut is reversed (i.e., the slider arrangement is at the strut'sconnection to the static structure rather than at its connection to thedivergent flap).

In minimum throat area/radius conditions, the synchronization ring 62 isshifted to the rearmost extreme of its range of motion. During thetransition of the synchronization ring, there is associated telescoping(contraction as shown) of the external flap. The need to accommodate asufficient range of telescoping across the throat area range may, asnoted above, exceed a desired range of extensibility associated with themode shift. Thus the mode strut may still operate to restrict a range ofmovement of the divergent flap and external flap combination.

FIG. 1 further shows the upstream flap 22 as including a main flap 150and a liner member 152 mounted inboard thereof. The exemplary linermember 152 includes a panel 154 extending from an upstream end 156 to adownstream end 158. One or more brackets 160 mount the liner member 152to the associated main flap 150, holding the liner member spaced aparttherefrom by a gap 162. In operation, a cooling air flow 164 passesthrough the gap 162 from an inlet 166 to an outlet 168. The exemplaryliner panel 154 has a length that is a major portion of an overalllength of the convergent flap. For example, the length may be measuredas a longitudinal length L_(PB) in a max throat high mode or low modecondition wherein the flap is oriented close to longitudinal. In thiscondition, the overall flap length is designated L_(FB). ExemplaryL_(PB) is approximately 80% of L_(FB).

Similarly to the convergent flaps, FIG. 2 shows the convergent seals 32as including a main seal member 180 and a liner member 182. The linermember 182 may have a similar panel and bracket construction to theliner member 152. Its panel 184 may have lateral portions configured tointerfit and cooperate with lateral portions of the adjacent panels 154.

The combined liner members of the convergent flaps and convergent sealsthus forms an overall liner. The liner cooperates with the convergentflaps and seals to create an interrupted (e.g., by the bracket legs)annular channel. The channel carries cool fan air for discharge into thegas turbine nozzle hot gas stream. The liner shields the adjacentportions of the convergent flaps and seals from exhaust gas heating. Thedischarged air provides a film cooling effect over the exposed surfaceof the convergent nozzle (flaps and seals) and downstream along thethroat and divergent nozzle.

We have determined that discharging cooling air further upstream mayproduce cooling benefits (discussed below). Thus, in accordance with thepresent teachings, the liner panel length may effectively be shortened.FIG. 3 shows a modified convergent flap 222 extending from an upstreamend 223 to a downstream end 224. The convergent flap 222 includes a mainflap 226 and a relatively short liner member 228 at an upstream endthereof. The enlarged view of FIG. 4 shows the liner member 228 asincluding a panel 230 extending from an upstream end 232 to a downstreamend 234 and having an inboard surface 236 and an outboard surface 238.The liner member also includes a mounting bracket 240. The panel has alongitudinal projected length L_(P). An exemplary length L_(P) is aminor portion of a projected overall length L_(F) and a minor portion ofa nearly similar main flap length. In the exemplary implementation, themain flap length is substantially close to the unprojecteddistance/spacing (S) and projected longitudinal spacing between the axes502 and 503. The panel 230 is mounted to an inboard surface 242 of themain flap by the mounting bracket 240. In the exemplary configuration,the main flap carries a threaded stud 244 that passes through anaperture in the bracket 240 and is secured to the bracket by a nut 246.A retainer clip 248 mounted to the main flap captures an upstreamportion of the bracket 240 to further retain the liner member. A channel250 thus extends between the panel and the main flap from an upstreaminlet 252 to a downstream outlet 254 and passes an air flow 256.

FIG. 5 shows a convergent seal member 260 and a divergent seal member262 between adjacent pairs of convergent and divergent flaps. Theconvergent seal member 260 includes a main seal 264 and a similarlyforeshortened liner 266 mounted to the main seal 264 in similar fashionto the mounting of the liner members 228.

FIG. 6 shows further details of an exemplary liner member 228. Theexemplary panel 230 is formed as an assembly of a main panel element orliner sheet 270, a backing sheet 272, and a flow blocker 274 secured byrivets 276. The exemplary liner sheet 270 includes a generally flat,rectangular central portion 280 that extends to the downstream end 224.An arcuate upstream deflector 282 extends upstream from an upstream endof the central portion 280 to the panel upstream end 223. The backingsheet 272 extends along a forward portion of the central portion 280 andalong the deflector 282, protruding slightly beyond the sides thereof.Lateral portions 286 and 288 of the liner sheet 270 extend along lateraledges of the central portion 280 and are curled outboard. Whereas theexemplary backing sheet 272 lies flat against the liner sheet 270, theexemplary flow blocker 274 has a main body 290 extending outboard fromthe liner and backing sheets and a pair of mounting tabs 292 engagingthe outboard surface of the backing sheet and secured by a lateral twoof the rivets 276.

The exemplary bracket 240 has a central web 300 and lateral webs or legs302 and 304 extending outboard. The legs 302 and 304 have upstream endportions 306 extending beyond an upstream end 308 of the web 300 andreceived in slots 310 in the flow blocker 274 and slots 312 in thebacking sheet 272.

FIG. 6 further shows a recessed area 320 and a central aperture 322 inthe central portion 300 for receiving the mounting stud and bolt.Additionally, a pair of locating pins 330 (discussed below) are shown.In an exemplary sequence of manufacture, the liner sheet 270, backingsheet 272, flow blocker 274, and bracket 240 are made from sheet stock(e.g., by stamping and forming). The bracket legs are then welded to theoutboard surface of the liner sheet. The welded assembly may then bewelded with a protective coating (e.g., a silicide coating). The backingsheet and flow blocker are then assembled to the liner sheet with theirslots receiving the bracket legs. The backing sheet and flow blocker maybe pre-coated prior to assembly. Once assembled, the rivets may beapplied.

For improving alignment of the bracket with the main flap, the exemplaryembodiment utilizes the locating pins 330. These register with holes 332in the bracket and corresponding holes (not shown) in the main flap. Theholes 332 may be drilled into the bracket prior to coating (e.g., afterwelding). Exemplary materials for the liner member components are hightemperature alloys. In the exemplary embodiment, the liner sheet 270,bracket 240, rivets 276, and locating pins 330 are formed of niobium(Nb), the backing sheet 272 is formed of nickel-based superalloy 625,and the flow blocker 274 is formed of nickel-based superalloy 718. Theseal liner members 266 may be similarly manufactured.

FIGS. 7 and 8 show further details of the assembled liner member 228.The flow blocker main body 290 has a central recess 336 along itsoutboard/distal edge 338. The recess 336 may be provided for clearancerelative to the clip 48. The height of the body 290 is selected toprovide a desired degree of flow restriction for the cooling flow 256.The deflector 282 (FIG. 7) has laterally recessed edge portions 340beyond which end portions 342 of the backing sheet protrude. The endportions 342, themselves, have recessed terminal portions 344 whichrespectively interfit outboard of protruding end portions 350 (FIG. 9)of a backing sheet 352 of the seal liner member 266. FIG. 9 shows theseal liner member 266 as also including a liner sheet 360 having anupstream deflector 362, a mounting bracket 364, a flow blocker 366 (FIG.10) and rivets 368.

The exemplary liner members 266 include outwardly recessed lateralportions 370 and 372 defining rebated/recessed areas 374 and 376 shiftedoutboard of a central portion 378. The recessed areas 374 and 376respectively accommodate the lateral portions 288 and 286 to permitinterfitting of the respective panels of the liner members 228 and 266(FIG. 11). The exemplary recessed areas 374 and 376 are tapered toaccommodate the range of convergence of the nozzle convergent flaps.

The present nozzle may be engineered as a redesign of an existing nozzleor otherwise engineered for an existing environment (e.g., as a drop-inreplacement for an existing nozzle such as the nozzle of FIG. 1 or theconvergent flaps and seals thereof). For example, the illustrated nozzlemay be formed as a retrofit kit for a baseline nozzle such as the nozzleof FIG. 1. In an exemplary retrofit, the convergent main flaps and mainseals and convergent flap and seal liner members may be replaced. Themain flaps and main seals may be dimensionally similar to thecorresponding baseline components but adapted to include appropriatelypositioned mounting studs or other features for the associated linermembers. The liner members may be substantially foreshortened relativeto the corresponding baseline components.

FIG. 12 shows the engine main (core) exhaust flow 400 meeting thecooling flow 164, 256 at the liner downstream end 158, 234 for thebaseline and reengineered nozzles/liners. The cooling flow provides filmcooling along the nozzle surface. The cooling effectiveness of the filmis believed to be a function of: 1) the distance traveled from the liner(the liner discharge plane); and 2) the mass flux ratio between coolantgas and core gas at the discharge plane. The coolant gas mixes with thecore gas and heats up as it travels the length of the exposed convergentand divergent sections. This phenomena is designated as filmeffectiveness decay, and is dependent on the mass flux ratio levels atthe discharge plane. As is discussed below, the cut-back linerdischarges the cooling gas film in a more favorable location at aforward portion of the nozzle. At this forward location thecoolant-to-core gas mass flux ratio is optimized, mixing is reduced, andthe enhanced film effectiveness level offsets the increased length oftravel.

The effects of this cooling flow may be determined at a point 402 adistance X along the nozzle downstream of the liner. The point 402 maybe a location of particular criticality (e.g., a location of maximumtemperature or thermal erosion). The point may be determinedexperimentally, or simply by post-use observation of the engine. As theliner is cut back (exit shifted upstream) by a given distance, theX-value of the particular point will increase by that distance.

The flows 400 and 164 each have a density ρ and a velocity V. FIG. 13shows a model plot 412 of the product of this density and velocity (themass flux) for the flow 400. FIG. 13 also shows a model plot 410 of massflux for the flow 164. The domain extends from zero to the baselineliner length L_(PB). The ratio of mass flux of the coolant flow 164 tothe mass flux of the exhaust flow 400 is designated the blowing ratio M.

FIG. 14 shows a model plot 440 wherein the domain is the ratio of theX-value of the chosen point 402 for the redesigned (cut-back)liner/nozzle to that value X_(B) of the baseline nozzle. The range isthe ratio of the blowing ratio M of the redesigned liner/nozzle to theblowing ratio M_(B) of the baseline nozzle. A line 442 divides a firstregion wherein cooling at the point 402 is improved (e.g., surfacetemperature reduced) from a second region wherein such cooling isreduced. Thus improved cooling appears to be achieved by increasedcut-back.

The exemplary nozzles have variable throat area (although the presentteachings may also be applied to other nozzles). For a typical variablenozzle throat area configuration, there is a partial nesting overlap oflateral portions of the flap liner panels and seal liner panels. Thedegree of overlap varies inversely as a function of nozzle jet area. Theinterfitting overlap features (whether actually overlapping in a min.throat condition, apart in a max. throat condition, or in between) blockflow and induce turbulence, interfering with film cooling effectiveness.The cut-back may reduce the maximum degree of overlap and may reduce theextent of the lateral overlap features thereby reducing or minimizingthe amount of film disturbance generated by the overlap features.

In the baseline nozzle the cooling air flow is relatively insensitve tothroat condition. This is because the liner exit is near the throat andthe static pressure there is relatively constant. In the cut-backnozzle, at high nozzle jet areas (e.g., at or near the max. throatcondition) the core pressure at the liner exit is reduced relative to anintermediate design throat area. This reduction causes the fan ductsystem to flow more coolant air to the liner system than at theintermediate throat condition. This enhanced flow rate offsets theenhanced mixing (due to the higher core velocity air at the high areacondition relative to the intermediate area condition). Similarly, forlow nozzle jet areas, the core velocity is reduced, mixing is reduced,and film effectiveness enhanced. The enhanced film levels offsets thereduced flow rate because at low jet areas the liner exit pressure isincreased and less flow is discharged through the liner system.Therefore the flow and film effectiveness impacts counteract each other.Thus, as in the baseline, there may be substantial independence ofcooling effectiveness and of nozzle jet area.

Protection of the convergent flaps/seals, however, imposes constraintson the cut-back. The cutback exposes a greater portion of the convergentflaps/seals to exhaust heating. Line 450 of FIG. 14 shows an X/X_(B)ratio below which (i.e., to the right) the convergent flap heatingexceeds an imposed threshold (e.g., a maximum temperature greater than atarget maximum). In a redesign/reengineering process, this may determinethe chosen liner length.

One or more embodiments of the present invention have been described.Nevertheless, it will be understood that various modifications may bemade without departing from the spirit and scope of the invention. Forexample, when implemented as a reengineering of an existing nozzle,various details of the existing nozzle may be preserved either bynecessity or for convenience. Additionally, the principles may beapplied to non-axisymmetric nozzles in addition to axisymmetric nozzlesand to vectoring nozzles in addition to non-vectoring nozzles.Accordingly, other embodiments are within the scope of the followingclaims.

1. A turbine engine nozzle subassembly comprising: an upstream flapassembly having a main flap and a liner, a cooling passageway formedbetween the main flap and liner; a downstream flap pivotally coupled tothe upstream flap for relative rotation about a hinge axis; and anactuator linkage coupled to at least one of the upstream flap and thedownstream flap for actuating the upstream and downstream flaps betweena plurality of throat area conditions, wherein: the liner has a trailingend spaced upstream from a trailing end of the main flap by at least 40%of a length of the main flap.
 2. The subassembly of claim 1 furthercomprising: an external flap pivotally coupled to the downstream flapand to an environmental structure so that a span between respectivecoupling locations with said downstream flap and environmental structureis extensible and contractable responsive to aerodynamic forces.
 3. Thesubassembly of claim 1 wherein the liner comprises: a liner body; and aliner mounting bracket secured to the liner body and to the main flap.4. The subassembly of claim 3 wherein: the liner body comprises anNb-based sheet and a Ni-based superalloy backing element.
 5. Thesubassembly of claim 1 wherein: the liner trailing end is spacedupstream from the main flap trailing end of the main flap by 70-80% ofthe length of the main flap.
 6. The subassembly of claim 1 wherein: theliner has a length of 15-50% of the length of the main flap.
 7. Thesubassembly of claim 1 wherein: the liner has a length of 20-30% of thelength of the main flap.
 8. A turbine engine nozzle comprising: a staticstructure; a plurality of flap subassemblies comprising: an upstreammain flap pivotally coupled to the static structure for relativerotation about an axis essentially fixed relative to the staticstructure; and a downstream flap pivotally coupled to the upstream flapfor relative rotation about a hinge axis; and a liner along the upstreammain flaps and forming a generally annular cooling air passageway, thecooling passageway having an outlet spaced upstream of a downstream endof the main flaps by a longitudinal distance of at least 40% of alongitudinal length of the upstream main flaps.
 9. The nozzle of claim 8wherein: the plurality of flap subassemblies are axisymmetricallyarranged about an engine centerline; said articulation is simultaneousfor each of the flap subassemblies; and each of the plurality of flapsubassemblies further comprises an external flap pivotally coupled tothe downstream flap.
 10. The nozzle of claim 8 wherein the linercomprises a circumferential array of: a plurality of first members, eachmounted to an associated one of the main flaps; and a plurality ofsecond members, each between an associated pair of the first members andmounted to an associated convergent seal.
 11. A turbine engine nozzlecomprising: a static structure; a convergent section comprising: acircumferential array of first flaps, each pivotally coupled to thestatic structure; a circumferential array of first seals, alternatinglyinterspersed with the first flaps; and a liner assembly; a divergentsection comprising: a circumferential array of second flaps, eachpivotally coupled to an associated one of the first flaps; and acircumferential array of second seals, alternatingly interspersed withthe second flaps, wherein the liner has an outlet spaced upstream of adownstream end of the main flaps by a longitudinal distance ofessentially at least 40% of a longitudinal length of the convergentsection.
 12. A gas turbine engine nozzle convergent section liner membercomprising: a panel having: an inboard surface; an outboard surface; aleading end; a trailing end first and second lateral ends; a lengthbetween the leading end and the trailing end; and a lateral span betweenthe first and second lateral ends, wherein: the lateral span is greaterthan the length.
 13. The liner member of claim 12 wherein: the length is40-60% of the lateral span.
 14. The liner member of claim 12 wherein:the panel has: a generally planar central portion; and means along thefirst and second lateral edges for interfitting with complementaryfeatures of a complementary panel.
 15. The liner member of claim 12further comprising: a mounting bracket secured to the panel andextending from the outboard surface and having: a central webessentially parallel and spaced apart from a central portion of thepanel and having a bolting aperture; and first and second lateral websextending toward the panel from first and second edges of the centralweb.
 16. The liner member of claim 12 wherein: the panel comprises aliner sheet, a backing sheet along only an upstream portion of the linersheet, and a deflector; a plurality of rivets securing the liner sheet,backing sheet, and deflector; and a pair of welds secure the mountingbracket to the liner sheet.
 17. A method for retrofitting a turbineengine or reengineering a turbine engine configuration which engine orconfiguration has or has previously had a first nozzle subassemblyhaving a convergent flap, a divergent flap, an external flap, and anactuation linkage coupled to the convergent flap, the method comprising:replacing a first liner member of the convergent flap with a secondliner member, the second liner member having a downstream end positionedupstream from a former position of a downstream end of the first linermember by at least 10% of a length of the convergent flap.
 18. Themethod of claim 17 wherein: said second liner member provides a highercoolant-to-gas ρv ratio than was provided by the first liner member. 19.The method of claim 17 wherein: said second liner member comprises aliner sheet and a mounting bracket welded to the liner sheet.
 20. Themethod of claim 17 wherein: a plurality of such first liner members of acircumferential array of such first nozzle subassemblies are replacedwith a plurality of such second liner members.